This invention relates generally to jet propulsion engines, and more specifically to fans having multistage cascade of outer airfoils mounted on fan blades.
An aircraft gas turbine engine typically includes a fan and a compressor. The compressor provides pressurized air to a combustor wherein the air is mixed with fuel and ignited for generating hot-high-pressure combustion gases. These gases flow downstream to one or more turbines that extract energy therefrom to power the fan and compressor and further expand to provide useful thrust for powering an aircraft in flight. In a turbofan engine, which typically includes a fan placed at the front of the core engine, a high pressure turbine powers the compressor of the core engine. A low pressure turbine is disposed downstream from the high pressure turbine for powering the fan.
In the art, some known gas turbine engine configurations incorporate an aft fan stage integral with a turbine rotor that is located near the aft end of the engine. In the art, fan configurations on a blade are referred to as a Fan on Blade (“FLADE™”). In a known blade/FLADE™, a radially inner portion of the blade works as a turbine blade that extracts energy from a flow of a hot combustion gas stream and a radially outer portion of the blade (FLADE™) works as a fan to impart energy into a flow of a relatively cooler air stream to raise its pressure. At least some known configurations of a FLADE™ on a turbine blade require an additional turbine stage and/or an additional turbine spool and related bearings, sumps, and other support structures. At least in some known applications, when the FLADE™ is located near the aft end of the engine, the axial distance available to collect and redistribute the combined flow may not be adequate and may have a high total pressure loss in the FLADE™ flow. The FLADE™ mounting on an aft located turbine may present difficulties in the installation of the engine into an aircraft if the FLADE™ flow needs to be directed from the engine to a site remote from the engine. A single stage FLADE™ may not be adequate to provide the required increase in pressure ratios in some applications. Also, a well known turbine design parameter, AN^2, where A is the annulus exit area and N is the rotational speed, may be made larger by the addition of a single-stage FLADE™ area. A large increase of the AN^2 parameter may require an extra-ordinarily heavy stage weight.
Accordingly, it would be desirable to have a multi-stage FLADE™ mounted on a fan stage located near the forward portion of a gas turbine engine. It would be desirable to have a fan FLADE™ that is capable of providing increased pressure ratio. It would be desirable to have a fan FLADE™ mounting that enhances the installation of the engine into an aircraft wherein the FLADE™ flow is capable of being directed to the engine exhaust nozzle or from the engine to a site remote from the engine.